System, method and apparatus for fabricating composite structures

ABSTRACT

Composite structures are fabricated and cured without applying external heat from an autoclave or oven. This technique uses composite preforms to form high strength, three-dimensional co-bonded joints, and thermoplastic conformal mandrels such as rota-molded vacuum bags. The bags form internal tooling within the structure and provide molding with integral fittings that circulate heated air within the bags. The exteriors of the bags are simultaneously under vacuum and exert pressure on the composite elements of the structure that are being cured. This internally-applied heat initially causes to the bags to soften and fully conform to the internal shapes of the structure that are being co-bonded. The heat then transfers into the uncured materials, causing them to cure without unnecessarily heating the entire structure or any required tooling.

BACKGROUND OF THE INVENTION

1. Technical Field

The present invention relates in general to fabricating compositematerials and, in particular, to an improved system, method andapparatus for fabricating composite structures without applying externalheat from an autoclave or oven.

2. Description of the Related Art

Composite materials comprising a fiber-reinforced resin matrix are oftenused to fabricate lightweight, high strength parts such as, for example,fuselages and stabilizers for aircraft. Optimized designs for thesetypes of parts will typically involve geometries that cause the internaltooling to become trapped upon cure. Fabrication of composite parts withtrapped geometries typically includes use of a solid mandrel having anexterior shape generally conforming to the desired interior shape of thepart. Uncured composite materials are then laid up on the mandrel andcured by applying heat and pressure according to well-known methods,

To address the problem of mandrel removal, the industry typically useseither segmented metal mandrels or expendable mandrels. Reusablesegmented metal mandrels have been used with a wide variety ofgeometries and sizes. They are extracted from a cured structure bydisassembling the mandrel and removing it piece by piece. These mandrelsare very costly and present handling problems due to their extremeweight and complexity, and because they are somewhat damage-prone.Expendable mandrels are typically made of plaster, water solubleeutectic salts or even eutectic metals. In the case of plaster, it isremoved by breaking it away from the part using impact devices after thecomposite part has been cured in an oven or autoclave. The brokenplaster pieces are then discarded at a significant cost to themanufacturer. The use of breakaway plaster is labor intensive, canresult in damage to cured composite parts, and produces large quantitiesof waste which are costly to dispose of.

The use of eutectic salts or metals may be environmentally hazardous,and although some percentage may be recoverable, recovery is not costeffective in many cases due to contamination and/or degradation of thematerial. These materials also tend to fuse to, and thereforecontaminate the interior surface of a structure, making it necessary toprovide a reliable barrier, which also needs to be subsequently removed.These mandrels are heavy, particularly those made with eutectic metals.Salt mandrels are comparatively fragile, therefore the handling of heavyand/or fragile mandrels presents yet another drawback.

All four types of mandrels are unsuitable for co-bonded structuresbecause they do not allow direct application of outward pressure to theinterior surfaces. Instead, pressure is applied only to the outside,requiring the outer surfaces and tooling to move inward. While thismethod does provide accurate interior mold lines, it is difficult tocontrol outer mold line contours and it becomes equally challenging tomaintain distortion-free fiber alignment. The substantial thermal massassociated with these materials presents yet another complication whenthe time comes to quickly and uniformly heat the structure during cure.This heat-up lag can increase the risk of having to scrap a structure bysimply forcing the cure process to go beyond its acceptable, specifiedparameters.

Expandable elastomeric soft tooling has occasionally been used as partof the bagging envelope in specific types of composite structurefabrication. In these cases, it is used to apply outward pressure tointerior surfaces. One such process utilizes an expandable tool wherepressure is applied to the interior of the tooling and expanded to forcethe uncured composite material to conform and consolidate against thesurfaces of an external tool. These mandrels are limited to simpledesign configurations in which dimensional control of the part'sinternal surfaces part is not critical. While this system does addressthe problem of consolidating internal elements of a composite structuresuch as those encountered in co-bonded assemblies, its flexible naturecannot provide a rigid backbone upon which these various elements can belocated or assembled prior to cure. Complex and cumbersome tooling musttherefore be used for that purpose.

As a result, non-recurring and recurring costs increase due to theadditional tooling required and/or the complications arising from havingto extract tooling prior to cure, or from having to bag around thesetooling elements. It is difficult for these mandrels to achieve uniformpressure distribution across the laminate during a cure cycle. Finally,elastomeric soft tooling is limited in its ability to be extractedintact from severely trapped areas. It is comparatively expensive,easily damaged and the elastomer itself cannot withstand more than a fewcure cycles before having to be replaced.

Traditional bagging systems, such as nylon film, also typically requiretooling to support the interior cavities. In addition, these systems aredifficult to conform to, and be extracted from, complex geometry. Thefilm has a tendency to bridge over corners, increasing the risk of bagfailure considerably during autoclave cure. This usually results in apoorly consolidated laminate which will have to be scrapped. Althoughtooling is typically added to minimize this risk, it becomes a laborintensive remedy.

U.S. Pat. No. 6,589,472, to Benson et al, discloses a method and toolingthat facilitates fabrication of trapped geometry composite structurewhile minimizing the aforementioned costs and risks. At roomtemperature, it acts as a rigid mandrel during lay-up and assembly of astructure. During cure, it is flexible and yet an inherently reliablevacuum bag. The tool provides support to the composite structure duringassembly, consolidates the laminate during cure, and is safely andeasily removable from the structure after it is cured.

That method of fabricating composite structure uses thermoplastics tocreate a conformal vacuum bag/temporary tool. The tool provides supportto elements within a composite structure prior to cure and improvedpressure distribution during cure by functioning as a high integritypressure intensifier/vacuum bag. The thermoplastic conformal tool/vacuumbag is designed to conform to the surfaces and cavities of the compositestructure, including closed or trapped geometry. The elements may beuncured when assembled and placed into the cure tool or fixture andexternally heated. Although the conformal tool becomes trapped in suchgeometry upon cure, it is removed by reheating the cured compositeassembly to an intermediate temperature. This softens the material tocollapse and allows it to be easily withdrawn.

Although that technique is workable for most composite components, somecomponents are far too large or cumbersome to be heated and cured incure-processing equipment (e.g., autoclaves or ovens). The manufactureof large, unitized complex composite structures, however, is highlydesirable from the standpoint of reducing part and fastener counts, andreducing weight. Unfortunately, the size of these structures istypically limited by the size of available ovens or autoclaves. In thecase of experimental programs, the lack of adequately-sized equipmentfrequently becomes the deciding factor as to whether a large-scale R&Dproject can be initiated. In a production mode, the cost of operatingautoclaves or other large equipment that are large enough to handleunusually large structures and their corresponding tooling also can be asignificant factor in the decision to proceed with the project.

A second limiting factor for such projects is the time required to heatand elevate the temperature of the tooling and structure to the requiredcure temperature at the desired heat-up rate. As the size of objectsincreases, the tooling required to hold the various structural elementsin position can become quite massive. In the autoclave or ovenenvironment that applies external heating, this tooling is typically thefirst mass to heat up, followed by the part itself. In the case ofco-bonded structures, this heating sequence is opposite of that which isdesirable. If the assembly is large enough, this heat-up lag canactually terminate the process prematurely since many of these materialshave a selected time-temperature “window” that must be met in order forthe properties of the material to be maximized.

A third limiting factor for these projects is the problem of out-time,or working life, of the uncured composite materials. If the requiredmaterials have a working life of less than two weeks, a point can bereached in which a large structure cannot realistically be assembled andprepared for cure within that span Thus, an improved solution thataddresses and resolves these limitations would be desirable.

SUMMARY OF THE INVENTION

Embodiments of a system, method, and apparatus for fabricating andcuring composite structures without the need for external heatingenvironments are disclosed. The invention utilizes several technologies,including woven composite preforms to form high strength joints, andthermoplastic conformal mandrels. The manufacture of large co-bondedstructures is enabled by the combined use of these technologies. Bytaking additional advantage of the unique characteristics of themandrels, this invention completely eliminates the need for a heatingenvironment that would otherwise apply heat externally to the toolingstructure.

In some embodiments, the composite parts being cured are locatedcompletely within the structure at the intersecting joints of theelements. The mandrels or bags may form internal tooling and vacuumbagging that may be molded with integral fittings. These fittings areused to attach ducting, through which heated air is circulated into theinterior of each individual bag. The exteriors of the bags aresimultaneously under vacuum and exert pressure on the elements beingcured. This exclusively internal heat source initially causes the bag tosoften and Fully conform to the shapes being co-bonded. The heat thentransfers to the uncured material, eventually causing it to cure.Although this technique may cause any needed large external tooling toeventually warm to some degree, the invention effectively removes itfrom the heat transfer cycle, which greatly improves heat transfer modeof the cure process.

In one embodiment, individual recirculating air heaters are linked to acommon controller to provide a tailored, non-fixed heat source forcuring the composite elements. The heating of structure is direct andzone-controlled, which offers the possibility of assembling and curing alarge structure in less cumbersome, more manageable segments.Significantly, this design helps achieve heat-up rates for compositeresin systems and removes concerns about non-uniform heat application.After the structure has been cured, the bags are removed. If they arewithin a trapped geometry of the structure, the same recirculating airheaters may be used to warm the bags to permit softening and easyremoval. The ducted heating system provides an ideal means of quicklyre-heating only the bags and removing them without subjecting the entirestructure to undesirable thermal stresses.

The foregoing and other objects and advantages of the present inventionwill be apparent to those skilled in the art, in view of the followingdetailed description of the present invention, taken in conjunction withthe appended claims and the accompanying drawings.

BRIEF DESCRIPTION OF THE DRAWINGS

So that the manner in which the features and advantages of the presentinvention are attained and can be understood in more detail, a moreparticular description of the invention briefly summarized above may behad by reference to the embodiments thereof that are illustrated in theappended drawings. However, the drawings illustrate only someembodiments of the invention and therefore are not to be consideredlimiting of its scope as the invention may admit to other equallyeffective embodiments.

FIG. 1 is a perspective view of a representative thermoplastic conformaltool/vacuum bag in accordance with this invention;

FIG. 2 illustrates the intended partial assembly of composite members.Actual assembly of these members will be accomplished using theconformal tool/vacuum bag of FIG. 1, as subsequently described;

FIG. 3 illustrates the assembly of composite members of FIG. 2 utilizingseveral thermoplastic members of FIG. 1 as locating tools or fixtures;

FIG. 4 illustrates a completed assembly of the composite members of FIG.3 and several conformal tools/vacuum bags of FIG. 1;

FIG. 5 is an enlarged cross-sectional view of a portion of thethermoplastic conformal tool/vacuum bag assembly of FIG. 4, taken alonglines 6-6 of FIG. 4;

FIG. 6 is an enlarged exploded, partial cross-sectional view of thethermoplastic conformal tool/vacuum bag assembly as shown in FIG. 5;

FIG. 7 is an enlarged cross-sectional view of a portion of thethermoplastic conformal tool/vacuum bag assembly as shown in FIG. 5,after it has been debulked and is ready to be placed in an oven orautoclave for curing;

FIG. 8 illustrates another embodiment of a thermoplastic conformaltool/vacuum bag in accordance with this invention. In this embodiment,the invention becomes an external vacuum bag with self-locatingfeatures, thereby facilitating the bagging of complicated geometries;

FIG. 9 is an exploded enlarged view of a portion of the thermoplasticconformal tool/vacuum bag of FIG. 8 taken along lines 10-10 of FIG. 8;

FIG. 10 is a schematic isometric view of one embodiment of a structuralmember being fabricated in accordance with the invention;

FIG. 11 is an enlarged isometric view of a portion of the structuralmember of FIG. 10, and is constructed in accordance with the invention;

FIG. 12 is a schematic isometric view of one embodiment of a compositecuring system constructed in accordance with the invention; and

FIG. 13 is a schematic isometric view of one embodiment of a heatingmodule for the composite curing system of FIG. 12 and is constructed inaccordance with the invention.

DETAILED DESCRIPTION OF THE INVENTION

Referring to FIGS. 1-13, embodiments of a system, method and apparatusfor fabricating large, unitized three-dimensional composite structureswithout an autoclave or oven are disclosed. The invention is well suitedfor various applications, such as those commonly encountered in themanufacture of aircraft and aerospace components.

The invention combines technologies involving composite preforms to formhigh strength joints, and thermoplastic conformal mandrels, and/or othermaterials and bag forming methods. For example, the invention maycombine three-dimensional, woven composite preforms to form highstrength three-dimensional co-bonded joints, and rota-molded vacuumbags, or RM bags. Other embodiments may comprise two-dimensional lay-upsfor various applications, such as corners. U.S. Pat. Nos. 6,589,472,6,676,882, 6,835,261, 6,849,150 and 6,712,099, are incorporated byreference in their entirety.

One component of the invention is directed to an improved method forfabricating a composite structure for, but not limited to, aircraft. Asused herein, the phrase “substantially conforms” is intended to mean ashape or geometry whose dimensions approximate a surface or cavity of acomposite structure or member. The phrases “tool/vacuum bag” and“mandrel” are used interchangeably.

Referring now to FIG. 1, a perspective view 10 of one embodiment of athermoplastic conformal tool/vacuum bag is shown. Thermoplasticconformal tool/vacuum bag 112 is a core member in this embodiment,structured to dimensions that substantially conform to a surface orcavity of a composite structure. Tool/vacuum bag 112 is rigid and hollowto provide support to at least a portion of the composite structure andto provide outward pressure distribution while the composite is beingcured. In this example, tool/vacuum bag 112 has a substantiallytriangular longitudinal cross-section and is generally wedge-shaped.Tool/vacuum bag 112 has thin rectangular opening 113 into its interiorto allow entry of recirculating heated air.

Tool/vacuum bag 112 may be prepared from conventional thermoplasticmaterials such as, but not limited to, acrylonitrile-butadiene-styrene(ABS) thermoplastic resin. Tool/vacuum bag 112 is fabricated viaconventional means, such as blow molding, vacuum-forming, rotationalmolding or the like.

As shown in FIG. 6, a liner 123 is wrapped around each tool/vacuum bag112 prior to its use as a locating tool. Liner 123 may include a film129 and a vent cloth 130. Film barrier 129, preferably TEFLON® film, isused to facilitate removal of tool/vacuum bag 112. Porous vent cloth 130is used to provide a pathway for volatile vapors, gases and trapped airto escape the assembly while the composite is being cured. Vent cloth130 may be a felt or fabric material.

For purposes of clarity, FIG. 2 shows a partial assembly of compositemembers only. During actual assembly and cure, the thermoplasticconformal tool/vacuum bag would be occupying the cavity 118. One or morecomposite skin elements 114, which are substantially flat plates, areplaced on a lay-up tool 119. A plurality of uncured composite jointmembers 124 are placed around the triangular, flat web members 128. Eachjoint member 124 is preferably formed of a woven fabric strip andcontains a resin matrix in or on joint member 124.

The assembled joint members 124/web members 128 are located onto therecessed edges of tool/vacuum bag 112. A plurality of assemblies(112/124/128) are placed onto skin element 114 at spaced intervalsdetermined principally by tool/vacuum bag 112, as shown in FIG. 3. Eachtool vacuum bag 112 is structured to dimensions that substantiallyconform to the intended cavity 118 as noted above. A second skin element114 is placed over the previously assembled elements to complete theuncured structure, as shown in FIG. 4. Skin 114 may be either cured oruncured, but in a preferred embodiment of the invention, it is alreadycured to provide the desired configuration and dimensional tolerances inorder to simplify the assembly tooling that is required. Also shown inFIG. 4, the upper and lower skins 114 do not join each other at thenarrow ends, leaving openings 113 at each cavity 118. Thecross-sectional geometry of cavity 118 depends on the particulargeometry of composite structure to be cured. In the embodiment of FIGS.1-7, the cross-sectional geometry of cavity 118 is substantiallyrectangular in shape.

Referring to FIGS, 5-7, composite joint member 124 may comprise anuncured composite whose shape resembles, but is not limited to, theGreek letter a or “pi” and has a longitudinal crossbar or base with twolongitudinal legs extending therefrom. A groove or channel 125 isdefined between the two legs. Web 128 is placed into channels 125 of theuncured composite joint member 124. The two legs of the uncuredcomposite joint member 124 closely receive and straddle the thickness ofthe web 128. Web 128 may be a cured or uncured composite material. Alsoweb 128 may be fabricated from metal.

The composite joint member 124 and web 128 are treated with a thermosetresin, such as an epoxy, to provide a bonding medium for these materialsduring cure. An overwrap 126 (FIG. 6) may be applied over compositejoint member 124 to be joined to improve bond strength between themember 124 and web 128. The overwrap is an uncured laminating materialsuch as a woven cloth or reinforcing fiber that may have Laminatingresin, such as epoxy, impregnated therein. Overwrap 126 maybe placed oncomposite joint member 124, extending from skin 114 to web 128.

Skins 114, composite joint members 124, webs 128, and tool/vacuum bags112 are assembled and sealed to an external vacuum bag which, in turn,is sealed to the external tool surfaces 119. A vacuum fitting piercesand seals into an appropriate portion of the external vacuum bag whichenvelopes the assembly. A vacuum hose is attached to the fitting andvacuum is drawn on the entire assembly. As shown in FIG. 7, the sides oftool/vacuum bag 112 expand outward, compressing composite members 124and pressing tightly against webs 128 and skins 114. This debulkingprocedure is well known to those skilled in the art.

While retaining the vacuum, the assembly shown in FIG. 7, is heatedinternally (i.e., within the individual bags 112), rather than applyingheat externally via an autoclave or oven, according to a thermal profilesuitable for curing the composite joint member 124. Structural bonds arethereby created that integrally link composite joint members 124 to webs128 and skins 114 to fabricate the desired composite structure.

As will be described herein greater detail, heat and pressure areapplied to the debulked assembly internally relative to the bags (i.e.,rather than externally with an oven or autoclave) according to atemperature and pressure profile appropriate for the thermosetting resinused. For example, in the case of epoxy laminating resins that are usedin most aerospace applications, the temperature for cure is generallyabout 350 degrees F. However, for the thermoplastic conformaltool/vacuum bag to soften and consolidate the laminate under pressure,the temperature should be at least about 30 to 50 degrees F. above theVICAT softening point (a standard ASTM test) of the particularthermoplastic material. For ABS thermoplastic resin, this ranges from210 to 320 degrees F. The vacuum pressure applied can range up toatmospheric (14 psia), depending on the resin system. The curing processcreates structural bonds that integrally link the composite to the webmembers and skin.

Following completion of the required cure cycle, the external vacuum bagand tooling, if any, are removed, yielding a completed cured assembly.The assembly is then internally re-heated (again, without an oven orautoclave) to a temperature below that which was reached during finalcure, but high enough to cause the conformal tools/vacuum bags 112 tore-soften and collapse. At this point, they are easily removed bypulling them outward through openings 113 (FIG. 4) and the thermoplasticmaterial is subsequently recycled or discarded. For ABS thermoplasticresin, the re-heat temperature may range from 285 to 320 degrees F., forsome embodiments.

The thermoplastic conformal tool/vacuum bag of this invention also canserve as a locating tool in most applications. FIG. 8 is an isometricview of another embodiment of the thermoplastic conformal tool/vacuumbag 142. The geometric form of tool/vacuum bag 142 provides the functionof locating and/or orienting the composite sub-elements during lay-upand cure, particularly for highly contoured skin members or hatstiffeners (not shown). Tool/vacuum bag 142 has a base portion 144 and aplurality of hollow channels 146 extending outwardly from base portion144. Base portion 144 has a straight section and a curved section,

As shown in FIG. 9, a blade stiffener 150 is placed within a channel 158of a composite joint member 156 that is similar to joint member 124 ofFIGS. 1-7. Joint member 156 may be treated with a thermosettable resin.An overwrap 152 may be placed over at least a portion of each blade 150and composite joint member 156. These sub-assemblies are placed into thechannels 146 of tool/vacuum bag 142 which is then placed and sealed ontoa cure tool. In doing this, tool/vacuum bag 142 simplifies locationand/or orientation of the structures during assembly and then functionsas a vacuum bag for debulking and cure, as discussed herein.

In some embodiments, the bags may be formed from cross-linkedpolyethylene, and are relatively stiff at room temperature to supportand hold the components in place prior to cure. The bags are generallyshaped in complement to the interiors of the cavities in which they arelocated. At elevated temperatures, however, the bags soften and almostcompletely conform to the cavity interiors as they transition to arubber-like texture.

Referring now to FIGS. 10-13, embodiments of a system, method, andapparatus for fabricating and curing large, unitized three-dimensionalcomposite structures without the need for an autoclave or oven areshown. For example, FIGS. 10-12 depict an exemplary wing box 201 havingupper and lower wing skins 114. The wing skins 114 are separated andsupported by one or more truss-like spars 207 that extend longitudinallywith respect to the wing box 201, and a plurality of ribs 209 thatextend laterally with respect thereto. The open spaces between the skins114, spars 207, and ribs 209 form the bays or cavities 118 previouslydescribed. Joints between these components are provided by the uncuredcomposite joint members 124 (FIG. 12), as described herein.

Each cavity 118 provides a location for one of the bags 112 describedherein. In FIG. 10, only two of the bags 112 are shown in two of thecavities 118, for illustration purposes. However, when fabricating wingbox 201, every cavity 118 would contain its own bag 112 as describedherein with respect to other embodiments. As best shown in FIG. 12, eachbag 112 is provided with one or more heated air ports 221 for pumpingheated air into bag 112, and one or more air exit ports 223 forreleasing air from bag 118. Ports 221, 223 permit heated air to becirculated into and out of each bag 118.

In the embodiment shown, each bag 118 and set of ports 221, 223 also isprovided with its own, independent bay heating module 225 (FIGS. 12 and13). Module 225 may be mounted to a platform 227 with casters 229 forportability, and daisy-chained or centrally controlled 226 with othermodules 225 for compressed air 231 for air flow, power 233 for heat, andother control features. Optional connections 235 may be provided onmodule 225 for overhead air and power drops, thereby eliminating triphazards. In the illustrated embodiment, module 225 has a data logger237, a programmed logic controller (PLC) 239, thermocouple (TIC) jacks241, a transducer line 243 and a vacuum source line 245.

Module 225 also is provided with one or more hot air injectors 251 forheated air port(s) 221, and one or more return air sources 253 for airexit port(s) 223. These injectors and ports may be linked via hoses,tubing, or the like. Thus, in one embodiment, each bag 118 isindividually heated by its own module 225 and cycled to cure adjacentones of the uncured preforms as described herein. In addition, the toolor lay-up table 119 may be formed with internal ports 228 for directlyheating the tool 119 (if required) with one or more additional,independent modules 225 that are likewise centrally controlled.

The invention has significant advantages. The method of this inventionis designed to reduce tooling and production costs, while improvingreliability (i.e., fewer tool/vacuum bag failures). In this method,pressure distribution in the tool/vacuum bag is improved and bridging issignificantly reduced or eliminated during cure of composite structuresat 350 degrees F. at a vacuum. Additionally, this method allows forgreatly simplified removal of tool/vacuum bags in difficult geometry,i.e., inner trapped locations. The thermoplastic conformal tool/vacuumbag is rigid at room temperature and may be used to build or lay up thecomposite members, i.e., temporary tooling that supports parts ormaterials during fabrication until completely ready for cure. Thetool/vacuum bag of this invention may be thermally formed to any desiredshape based on the geometry of the composite structure. Thethermoplastic conformal tool/vacuum bag re-forms during theconsolidation process of cure and then collapses when re-heated for easyremoval from trapped geometries. It also enables the fabricator togreatly reduce the number of fasteners used in typical compositestructure assemblies and do so in a simplified manner.

For manufacturers of large-scale unitized composite structures, theinvention eliminates the need for expensive, fixed facilities such aslarge ovens and autoclaves. The size of a manufactured structure is nolonger a limiting factor. Instead of having to move the assembly to anoven or autoclave, it can now remain stationary. This reduces the riskand expense involved in transporting large assemblies. The cure processequipment now becomes portable as it can be easily taken to any worksite. This design greatly reduces the logisitics and complexity ofmanufacturing large structures.

The system further provides much more precision in controlling both theheat-up rate and the overall temperature of an assembly during cure,which enhances quality control. In addition, material out-time concernscan be eliminated by assembling and curing discrete sections of astructure, one at a time. Moreover, thermoplastic bag removal afterco-bond or cure is greatly simplified without inducing unwanted thermalstresses to the cured structure.

In one embodiment, the invention provides numerous advantages for thewing box application. The porting through spar perforations to the baginteriors provides quick and even heat transfer, and simplified ventingof air and volatiles to the vacuum source. The invention providesgreater dimensional control by applying pressure and heat toward outermold line (OML) surfaces and substructure locations. Mechanical supportof truss-like spars and rib webs obviates pressure differentials, andthere is a low risk of sealing one bag to another.

While the invention has been shown or described in only some of itsforms, it should be apparent to those skilled in the art that it is notso limited, but is susceptible to various changes without departing fromthe scope of the invention.

1. A method of fabricating and curing a composite structure, comprising:(a) providing a composite structure having uncured components, andcavities of open spaces defined between the uncured components; (b)positioning a bag in at least some of the cavities of open spaces; (c)providing each bag with inlet and outlet air ports; and (d) circulatingheated air through each bag via respective ones of the inlet and outletports to cure the uncured components.
 2. A method according to claim 1,further comprising cooling the bags through said respective ones of theinlet and outlet ports to cool the cured composite structure, and thenreheating the bags through said respective ones of the inlet and outletports to a temperature lower than a temperature of step (d) to softenthe bags and remove them from the cured composite structure.
 3. A methodaccording to claim 1, wherein step (d) comprises heating the bags with abay heating module.
 4. A method according to claim 3, wherein the bayheating module comprises a plurality of independent bay heating modulesthat are collectively controlled for air flow and heat.
 5. A methodaccording to claim 3, wherein the bay heating module comprises a datalogger, a programmed logic controller (PLC), thermocouple jacks, atransducer line, and a vacuum source line.
 6. A method according toclaim 3, wherein the bay heating module has at least one hot airinjector for the inlet port, and at least one return air source for theoutlet air port, which are linked via hoses.
 7. A method according toclaim 1, wherein step (d) occurs without an autoclave or oven.
 8. Amethod of fabricating and curing a composite structure, comprising: (a)providing a composite structure having cured components and uncuredcomponents, the uncured components forming joints for the curedcomponents, and cavities of open spaces are defined between the curedand uncured components; (b) positioning a bag in each of the cavities ofopen spaces; (c) providing each bag with inlet and outlet air ports; and(d) circulating heated air through each bag via respective ones of theinlet and outlet ports to cure the uncured components.
 9. A methodaccording to claim 8, further comprising cooling the bags through saidrespective ones of the inlet and outlet ports to cool the curedcomposite structure, and then reheating the bags through said respectiveones of the inlet and outlet ports to a temperature lower than atemperature of step (d) to soften the bags and remove them from thecured composite structure.
 10. A method according to claim 8, whereinstep (d) comprises heating the bags with a bay heating module.
 11. Amethod according to claim 10, wherein the bay heating module comprises aplurality of independent bay heating modules that are collectivelycontrolled for air flow and heat.
 12. A method according to claim 10,wherein the bay heating module comprises a data logger, a programmedlogic controller (PLC), thermocouple jacks, a transducer line, and avacuum source line.
 13. A method according to claim 10, wherein the bayheating module has at least one hot air injector for the inlet port, andat least one return air source for the outlet air port, which are linkedvia hoses.
 14. A method according to claim 8, wherein step (d) occurswithout an autoclave or oven.
 15. A method according to claim 8, whereinthe composite structure is a wing box, the cured components compriseupper and lower wing skins, spars and ribs, and the uncured componentscomprise three-dimensional woven composite preforms.
 16. A method offabricating and curing a composite structure, comprising, (a) providinga composite structure having cured components and uncured components,the uncured components forming joints for the cured components, andcavities of open spaces are defined between the cured and uncuredcomponents; (b) positioning a bag in each of the cavities of openspaces; (c) providing each bag with inlet and outlet air ports; (d)circulating heated air through each bag with a bay heating module viarespective ones of the inlet and outlet ports to cure the uncuredcomponents without an autoclave or oven; and (e) cooling the bagsthrough said respective ones of the inlet and outlet ports to cool thecured composite structure, and then reheating the bags through saidrespective ones of the inlet and outlet ports to a temperature lowerthan a temperature of step (d) to soften the bags and remove them fromthe cured composite structure.
 17. A method according to claim 16,wherein the bay heating module comprises a plurality of independent bayheating modules that are collectively controlled for air flow and heat.18. A method according to claim 16, wherein the bay heating modulecomprises a data logger, a programmed logic controller (PLC),thermocouple jacks, a transducer line, and a vacuum source line.
 19. Amethod according to claim 16, wherein the bay heating module has atleast one hot air injector for the inlet port, and at least one returnair source for the outlet air port, which are linked via hoses.
 20. Amethod according to claim 16, wherein the composite structure is a wingbox, the cured components comprise upper and lower wing skins, spars andribs, and the uncured components comprise three-dimensional wovencomposite preforms.